This application is related to a co-pending application entitled xe2x80x9cPermanent Magnet Phase-Control Motorxe2x80x9d, Ser. No. 09/624,554 filed concurrently herewith, which application is incorporated herein by reference in its entirety.
The present invention relates to an actuator for a rotor blade on a rotorcraft and, more particularly, to a hub mounted actuator for controlling the position of a rotor blade or a flap attached thereon.
A rotor system on a rotorcraft includes a rotor head mounted on a rotating shaft. The rotor head connects to rotor blades that are individually pivotable about a feathering axis to control the aircraft flight. In conventional rotorcraft, the pivoting or pitch of the blade, as well as the orientation of the rotor tip-path plane (the plane of the rotor blade tips)with respect to the shaft, are controlled using a mechanical linkage arrangement that typically includes a swashplate which is connected to the rotor blades with control or push rods. The swashplate includes a rotating portion that supports the control rods and is linked to the rotor shaft through a scissor linkage. As the swashplate is a rigid or near rigid member, the motion that can be transmitted to the blades is limited. The swashplate also includes a non-rotating portion which is attached to the rotating portion through a bearing. The non-rotating swashplate is moved by the primary control actuators, to simultaneously vary the pitch of all of the rotor blades. For conventional helicopters, the primary control actuators are low bandwidth and, thus, actuation is limited to steady or once per rev (1P) on the rotor. Inclusion of higher bandwidth primary actuators expands this range, but there remains a physical limitation that the rotor blades can only be practically actuated at Nxe2x88x921, N and N+1 per rev, where N is the number of rotor blades resulting from the primary actuators being actuated at N per rev. Actuation at these higher frequencies has been accomplished and is referred to as Higher Harmonic Control (HHC).
There are times, however, when it is desirable to control blade motions at a frequency other than the frequencies allowed by conventional trim control or HHC control. The capability of controlling rotor motion at multiple and arbitrary frequencies can yield significant benefits in rotor performance and load factor capability. Rotor blade control at these frequencies can also be used to reduce aircraft vibrations and externally radiated noise.
For example, helicopter main rotor lift and rotor driving torque produce reaction forces and moments on the helicopter main gearbox. In addition to these primary flight loads, the aircraft is also subjected to vibratory loads originating from the main rotor system. These vibratory loads produce vibrations within the aircraft that are extremely bothersome and fatiguing to the passengers. The vibratory loads on the main rotor are generated from the passage of the rotor blades through their own complex wake structure. Unlike a fixed wing aircraft where the wing wake trails harmlessly behind the aircraft, a helicopter wing or blade, due to the rotation of the rotor, must pass through its own wake repeatedly and in a manner dependent on, among other things, the weight of the helicopter and its forward speed. Unsteadiness and spatial variations in the incident air flows yield vibratory loads on the blade that, when integrated into the rotor hub, result in vibratory loads which are then transmitted into the airframe cockpit and cabin. The frequency content of these vibratory loads is dominated by multiples of the number of blades (i. e. for a 4-bladed helicopter, the principle frequency is 4 times the rotor speed-additional but smaller contributions are multiples of this, e.g. 8 times rotor speed, 12 times rotor speed, etc.)
In addition to vibration, the interaction of the rotor blades with blade vortices developed by the preceding blades during rotation generates external noise. As the rotor blade rotates, the changing distribution of lift along the blade and over time results in trailed and shed vortices (concentrated rotational air flow). During normal flight modes, these blade vortices do not cause any particular problem. However, in certain instances, for example when the aircraft is descending in an approach, the following blades come into proximity to these blade vortices generating an impulsive noise or slap. These blade-vortex interactions (BVI) produce an external noise signature which can be easily detected at long range, increasing the aircraft""s vulnerability when in a hostile environment, and for commercial applications, becomes a community annoyance.
Many attempts have been made over the years to alleviate or reduce blade vortex interactions. A considerable amount of those attempts have been directed toward passive type systems wherein the blade is designed to weaken the vortex at the blade tip. See, for example, U.S. Pat. No. 4,324,530 which discloses a rotor blade with an anhedral swept tapered tip which reduces the intensity and shifts the location of the tip trailing edge vortex so as to reduce the occurrence of blade vortex interactions and improve blade performance.
While passive solutions have provided some reduction in blade vortex interaction, these types of solutions are near their limits in terms of noise reduction benefits. In order to meet more stringent military and commercial requirements, alternate solutions need to be considered.
Active rotor control systems have recently been proposed to counteract blade vortex interactions. These systems are typically designed to change the motion of the rotor blade and the trailed vortices to increase the distance between them so as to reduce the magnitude of the interaction and thus to reduce the level of generated noise. One of-these systems is called higher harmonic blade pitch control (HHC, as mentioned previously) wherein the blade pitch is controlled to maximize the miss distance between the blades and the trailed vortices. This type of system was originally developed to address rotor induced vibration. Results to date have shown success at reducing both vibration and noise, independently. The limitation in allowable rotor blade motions, however, has kept researchers from achieving benefits in both noise and vibration simultaneously. In contrast, results have shown that improved vibration solutions generally increase noise levels and vice-versa. Other major concerns are the high loads imparted to the blades due to inertial loads generated by moving the entire blade as well as the high levels of primary actuation power needed to control the blade motions. An additional problem is the potential for excessive wear occurring on the flight critical primary actuator seals.
Another active control system is discussed in U.S. Pat. No. 5,588,800. This active control system is mounted within a helicopter rotor blade and includes actuatable flaps on the rotor that are controlled to reduce the blade vortex interaction. An actuator is used to control the movement of the flaps and can be either mechanical, electrical, pneumatic, or hydraulic. While U.S. Pat. No. 5,588,800 states that the blades are actuated according to a prescribed schedule to reduce the development of BVI, there is no discussion in U.S. Pat. No. 5,588,800 about how such control is provided (i.e., the type of system used to actuate the flap).
U.S. Pat. No. 5,639,215 discloses a similar actuatable flap assembly. In this assembly, the actuator is a mechanical actuator that is either a push-rod type device, a linkage, or a servo-motor driven rack. As with U.S. Pat. Nos. 5,588,800, 5,639,215 does not disclose the type of system is used to control the actuator.
Alternate types of systems are currently being investigated for actively controlling a rotor blade, including piezoelectric patches, piezofiber composites, PMN stacks (i.e., a series of piezo-electric wafers stacked up with varying orientation), and magneto-strictive stacks (i.e., stacks of Terfenol rods, for example, encased in magnetic coils which produce a magnetic force that elongates the Terfenol rods). These devices (also known as smart materials) are extremely displacement limited and thus require complex displacement magnification schemes. In sum, this results in a very heavy solution.
Although the prior art systems for actively controlling the rotor blade interactions with the blade vortex are empirically better than passive systems, these prior art systems do not adequately address the realistic problem associated with controlling a flap or rotor blade at a frequency other than the rotational speed of the rotor shaft.
Another issue with systems for controlling blade actuation is the need to transmit power to the rotating blade actuation system. For example, if a hydraulic, pneumatic or electrical actuation system is mounted on the rotating rotor hub, then power must be transmitted from the aircraft through some means, (e.g., a slip ring) to the rotating hub. Slip rings, however are a wear item and suffer reliability and maintainability issues. In addition, the size of such devices make integration into the rotor hub and transmission design quite difficult.
One prior art system for controlling the actuation of a rotor blade was developed by ZF Luftfahrttechnik, and EuroCopter Deutschland and tested on the BO-105 helicopter. In this system, the conventional blade pitch control rods were replaced with servo-actuators which allowed the pitch of each blade to be independently controlled in a range of between 2/rev to 6/rev. There are several drawbacks to this system. First, moving the blade at the root requires a high force and, thus, is not efficient. Also, the servo-actuator in this system is a flight critical component and requires an auxiliary hydraulic pump with a large flow rate to accommodate the high force required. In addition, the need for a hydraulic slip ring brings both weight, complexity and reliability and maintainability issues to the system.
A need, therefore, exists for an improved actively controlled blade actuation system for controlling blade pitch or flap changes.
The present invention relates to a hub mounted actuation system for providing control of a portion of a rotor blade, such as a flap, on a rotorcraft. The rotor blade is attached to a rotor hub and shaft that rotates with respect to an airframe. The hub mounted actuation system includes a stationary support mounted to the airframe and a rotary support attached to the rotor hub for concomitant rotation with the rotor shaft.
At least one hub actuator rotates in combination with the rotor blade and includes a piston which is slidable within a housing. The piston and housing define a pressure chamber within the actuator which contains a fluid to be pressurized.
Displacement control means is disposed between the stationary support and the rotary support for controlling movement of the piston within the housing. A portion of the displacement control means is attached to the stationary support and a portion is attached to the hub actuator.
A linkage connects the hub actuator to the portion of the blade to be controlled. The linkage is adapted to displace the blade portion as a function of the movement of the piston within the housing.
In one embodiment of the invention, the displacement control means includes a stator mounted to the stationary support, and an intermediate member supported by and rotatable relative to the rotary support. One of either the piston or the housing is attached to the intermediate member and the other is attached to the rotary support. The intermediate member has a plurality of magnets spaced around its circumference adjacent to the stator on the stationary support such that when current is supplied to the stator a magnetic field is created that causes the intermediate member to rotate relative to the rotary support. Rotation of the intermediate member relative to the rotary support causes the piston to move within the housing.
In another embodiment of the invention, the displacement control means includes a deformable cam mounted between the stationary support and the rotary support. The deformable cam is supported by cam actuators mounted to the stationary support. The cam actuators are adapted to deform the cam as commanded. A cam follower is in rolling contact with the deformable cam and attached to the piston of the hub actuator. As the cam follower rolls around the cam, the deformation of the cam causes the piston to move within the housing of the hub actuator.
The foregoing and other features and advantages of the present invention will become more apparent in light of the following detailed description of the preferred embodiments thereof, as illustrated in the accompanying figures.